The field of the present invention relates generally to gas turbine engines and, more particularly, to turbine engine rotor blades and a method of assembling a turbine rotor blade assembly.
FIG. 1 is a perspective view of a pair of known rotor blades that each include an airfoil 2, a platform 4, and a shank or dovetail 6. During fabrication, the known rotor blades are cast such that the platform is formed integrally with the airfoil and the shank. More specifically, the airfoil, the platform, and the shank are cast as a single unitary component.
During operation, because the airfoil is exposed to higher temperatures than the dovetail, temperature gradients may develop at the interface between the airfoil and the platform, and/or between the shank and the platform. Over time, thermal strain generated by such temperature gradients may induce compressive thermal stresses to the platform. Accordingly, the increased operating temperature of the platform may cause platform oxidation, platform cracking, and/or platform creep deflection, which may shorten the useful life of the rotor blade.
To facilitate reducing the effects of the high temperatures in the platform region, shank cavity air and/or a mixture of blade cooling air and shank cavity air is introduced into a region below the platform region using cooling passages to facilitate cooling the platform. However, the cooling passages may introduce a thermal gradient into the platform which may cause compressed stresses to occur on the upper surface of the platform region. Moreover, because the platform cooling holes are not accessible to each region of the platform, the cooling air may not be uniformly directed to all regions of the platform.
Since the platform is formed integrally with the dovetail and the shank, any damage that occurs to the platform generally results in the entire rotor blade being discarded, thus increasing the overall maintenance costs of the gas turbine engine.